US11408295B2 - Nozzle for a turbine, turbomachine turbine equipped with said nozzle and turbomachine equipped with said turbine - Google Patents

Nozzle for a turbine, turbomachine turbine equipped with said nozzle and turbomachine equipped with said turbine Download PDF

Info

Publication number
US11408295B2
US11408295B2 US17/419,404 US201917419404A US11408295B2 US 11408295 B2 US11408295 B2 US 11408295B2 US 201917419404 A US201917419404 A US 201917419404A US 11408295 B2 US11408295 B2 US 11408295B2
Authority
US
United States
Prior art keywords
nozzle
sector
angular
tab
foot
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
US17/419,404
Other languages
English (en)
Other versions
US20220074312A1 (en
Inventor
Coralie Cinthia Guerard
Hamza Guessine
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Safran Aircraft Engines SAS
Original Assignee
Safran Aircraft Engines SAS
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Safran Aircraft Engines SAS filed Critical Safran Aircraft Engines SAS
Assigned to SAFRAN AIRCRAFT ENGINES reassignment SAFRAN AIRCRAFT ENGINES ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: GUESSINE, HAMZA, GUERARD, CORALIE CINTHIA
Publication of US20220074312A1 publication Critical patent/US20220074312A1/en
Application granted granted Critical
Publication of US11408295B2 publication Critical patent/US11408295B2/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material
    • F01D11/125—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material with a reinforcing structure
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F05D2220/323—Application in turbines in gas turbines for aircraft propulsion, e.g. jet engines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/128—Nozzles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/55—Seals
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/30—Retaining components in desired mutual position
    • F05D2260/33—Retaining components in desired mutual position with a bayonet coupling

Definitions

  • the present invention relates to a nozzle for a low- or high-pressure turbine, a low- or high-pressure turbine comprising this nozzle and a turbomachine such as an aircraft turbojet or turboprop comprising the above turbine.
  • Some turbomachines comprise a low-pressure turbine and a high-pressure turbine, which recover some of the energy released by the combustion of fuel to drive a low-pressure compressor and a high-pressure compressor respectively, located in the region of the air intake of the turbomachine.
  • FIG. 1 shows a multi-stage 10 low-pressure turbine 1 , each stage 10 comprising successively a nozzle 2 and a rotor wheel 3 , from upstream to downstream relative to the direction of flow of air in the turbine (arrow V: that is, from left to right in FIG. 1 ). All rotors are fixed on a shaft, not shown in the figures and driven in rotation simultaneously.
  • the low-pressure turbine 1 has a longitudinal axis X-X′.
  • Each nozzle 2 is sectorised, that is, formed from several angular nozzle sectors 20 , arranged circumferentially end-to-end and assembled together.
  • Each angular nozzle sector 20 comprises two platforms in a portion of an arc of a circle, arranged coaxially one inside the other and connected together by radial or substantially radial blades 201 . These platforms delimit between them the annular duct for gas flow in the turbine 1 .
  • the external platform 202 comprises means for hooking onto an external casing 100 of the turbine 1 .
  • the internal platform 203 extends towards the interior of the turbine by a radial foot 204 which terminates in a baseplate 205 .
  • this baseplate 205 On its radially internal face this baseplate 205 bears elements made of abradable material 206 .
  • This abradable material is in the form of a honeycomb structure.
  • the elements made of abradable material 206 cooperate with annular sealing lips 30 carried by the adjacent rotor 3 to form sealing joints of “labyrinthine seal” type.
  • the sealing lips 30 sink furrows into the abradable material 206 .
  • the honeycomb 206 is brazed onto the baseplate 205 of each angular nozzle sector 20 , then is renewed by machining.
  • this material is machined by electro-erosion (known by the acronym EDM for “Electrical Discharge Machining”), which is an expensive technique requiring specific and expensive controls.
  • each sector 20 has sealing slots 207 which terminate on each of the end faces of a sector 20 intended to make contact with the end faces of the adjacent sector of nozzle 20 .
  • An inter-sector sealing tab (not shown in the figures) is engaged in each sealing slot 207 so as to overlap the space in between two adjacent sectors of nozzle 20 and stop or limit inter-sector air leaks.
  • the aim of the invention is to propose a turbomachine nozzle which resolves the above disadvantage of the prior art.
  • the aim of the invention is to provide a turbomachine nozzle which is simple and less costly to produce and assemble while decreasing air leaks in the duct.
  • Another aim is to simplify maintenance of the nozzle.
  • the invention relates to a nozzle for a turbine comprising several angular nozzle sectors, each angular sector comprising two sectors of platforms respectively internal and external, in the form of an arc of a circle, coaxial, connected together by several radial or substantially radial blades, each sector of internal platform being attached to a radially internal foot, this nozzle comprising an annular collar to which said angular nozzle sectors are fixed end-to-end circumferentially, this collar comprising a cylindrical ring of which the radially internal face bears an abradable material.
  • the radially internal foot of each sector of internal platform comprises a tab which extends radially towards the interior from said foot, and which extends circumferentially over part of the length of said foot, the radially external face of said cylindrical ring bearing an external radial wing fitted with a plurality of L-shaped legs of which one of the arms called “fastening” is attached to the external radial wing and extends perpendicularly to a radius of the ring and parallel to the longitudinal axis X 1 -X′ 1 of said ring and the other arm of which, called “retaining”, extends opposite this external radial wing and at minimal distance from the latter so as to form a slot with it for receiving said tab of each angular nozzle sector, and this slot is open at one of its longitudinal ends to allow engagement of said tab to ensure fastening by coupling of said collar on each angular nozzle sector.
  • the invention also relates to a turbomachine turbine comprising the above nozzle and a turbomachine, such as an aircraft turbojet or turboprop, which comprises the above turbine.
  • FIG. 1 is a view in longitudinal section of a low-pressure turbine of the prior art
  • FIG. 2 is a perspective view of the collar of the nozzle according to the invention.
  • FIG. 3 is a detailed view of FIG. 2 .
  • FIG. 4 is a detailed view of FIG. 3 .
  • FIG. 5 is a perspective view of an angular nozzle sector according to the invention.
  • FIG. 6 is a perspective view of a portion of collar and of an angular nozzle sector assembled according to the invention
  • FIG. 7 is a schematic view in axial section of the collar, of the internal part of an angular nozzle sector and a shielding plate joined together, the plate being formed according to a first embodiment
  • FIG. 8 is a schematic view in axial section of the collar, the internal part of an angular nozzle sector and a shielding plate joined together, the plate being formed according to a second embodiment
  • FIG. 9 is a perspective view of an angular nozzle sector, two shielding plates and part of the collar, assembled, which constitute part of a nozzle according to the invention.
  • FIG. 10 is a detailed and perspective view of two different shielding plates joined together respectively on the collar and on an angular nozzle sector,
  • FIG. 11 is a detailed and perspective view of a variant embodiment of a shielding plate assembled on an angular nozzle sector.
  • FIG. 12 is a detailed and perspective view of a fastening mode of an angular nozzle sector on the collar.
  • the nozzle according to the invention is referenced 4 and comprises an annular collar 5 , on which several angular nozzle sectors 6 are fixed. Only part of the nozzle 4 is visible in FIG. 9 .
  • the annular collar 5 will now be described in more detail in conjunction with FIGS. 2 to 4 . It extends over 360° and has a longitudinal axis X 1 -X′ 1 .
  • the annular collar 5 comprises a cylindrical ring 50 (or tube) of minimal width relative to its diameter, of minimal thickness and having a longitudinal axis X 1 -X′ 1 .
  • This ring 50 has a radially internal face 51 and a radially external face 52 .
  • the internal face 51 bears an abradable material 53 made in one or more elements.
  • This abradable material 53 is preferably a honeycomb structure. It is brazed onto the ring 50 and renewed by machining as described earlier for the abradable material 206 .
  • This abradable material 53 is also cylindrical.
  • the external face 52 of the ring 50 comprises several first mechanical fastening members, distributed over its periphery and intended to cooperate with second complementary mechanical fastening members carried by each of the angular sectors 6 .
  • these mechanical fastening members are of “coupling” type and enable the collar 5 and the nozzle sectors 6 to be coupled together by engaging of the tabs in slots for receiving these tabs.
  • FIGS. 3 and 4 show that the collar 5 is fitted with a radial wing 54 which extends radially towards the exterior from the external face 52 of the ring 50 .
  • This wing 54 is annular and extends over the entire circumference of the ring 50 . It has a rear face 541 and an opposite front face 542 .
  • the rear face 541 is preferably oriented to downstream of the turbine.
  • a plurality of L-shaped legs 55 is attached to the wing 54 and distributed uniformly over the full circumference of the latter.
  • Each leg 55 is also incurved and has a general form of an arc of a circle of which the circle is coaxial to the ring 50 , but has a radius larger than the radius of the ring 50 .
  • the leg 55 fits the shape of the ring 50 .
  • Each L-shaped leg 55 has a short arm 551 called “fastening arm” and a long arm 552 called “retaining arm”.
  • the fastening arm 551 extends perpendicularly to a radius of the ring 50 and parallel to the longitudinal axis X 1 -X′ 1 of said ring.
  • the retaining arm 552 extends circumferentially and at minimal distance from the rear face 541 of the wing 54 so as to form a slot 553 with the latter.
  • the long arm 552 has a front face 554 which extends opposite and parallel to the rear face 541 of the wing 54 .
  • a part of the arm 551 constitutes the bottom 555 of the slot 553 .
  • the slot 553 is open opposite the bottom 555 , between the free end of the long arm 552 and the wing 54 .
  • the long arm 552 has a lug 556 (see FIG. 4 ) which projects from the front face 554 in the direction of the rear face 541 of the wing 54 so as to block the slot 555 partially only.
  • the arm 552 is advantageously slightly flexible.
  • the ring 50 can optionally be split (see the longitudinal slot 500 ). This lends it some flexibility and makes it easier to mount the nozzle 4 .
  • the different parts 50 , 54 and 55 of the collar 5 are preferably monobloc, (made in a single piece).
  • the collar 5 can be made by machining or by additive manufacturing, for example.
  • the collar 5 is preferably made of metal.
  • FIGS. 5 to 8 An example of embodiment of an angular nozzle sector 6 will now be described in conjunction with FIGS. 5 to 8 .
  • this sector 6 comprises two angular sectors of platforms in an arc of a circle, coaxial, specifically a sector of external platform 61 and a sector of internal platform 62 , connected together by several radial or substantially radial blades 63 .
  • the different angular sectors 6 are assembled together around the collar 5 , end-to-end circumferentially, so that the different angular sectors of external platform 61 jointly form the external platform of the nozzle 4 and the different angular sectors of the internal platform 62 jointly form the internal platform of the nozzle 4 .
  • the angular sector of internal platform 62 is attached to a foot 621 , called “internal radial foot”, as it extends radially from the internal face of said sector of internal platform in the direction of the interior of the nozzle 4 .
  • the internal radial foot 621 is incurved, viewed front on. It is extended radially towards the interior by a radial internal tab 622 , also incurved, viewed front on.
  • the tab 622 is centred in length relative to the internal radial foot 621 so as to avoid a cantilever.
  • the length of the tab 622 is preferably reduced to what it needs to absorb forces.
  • the tab 622 has a length L 1 tangentially, equal to around one third of the length L 2 tangentially of the foot 621 .
  • the foot 621 supports a single tab 622 . But according to the dimensions of the angular sectors of nozzle 6 , it is possible to have several tabs 622 on each foot 621 .
  • the tab 622 is not as thick as the foot 621 so that with this foot 621 it can form either a shoulder 623 which terminates on the front face 6211 of the foot 621 (see FIG. 7 ), or a shoulder 624 which terminates on the rear face 6212 of the foot 621 (see FIG. 8 ).
  • one or more sealing plates 7 can be fixed to the different angular nozzle sectors 6 .
  • the plate 7 is formed from a profile slightly curved in an arc of a circle, according to a circle of which the center is coaxial to the one of the circle of the internal platform 62 in an arc of a circle so as to mould to the form of this platform.
  • the plate 7 has a substantially L-shaped form, with a fastening flank 71 and a shielding flank 72 .
  • the fastening flank 71 is intended to be fixed to the foot 621 , for example by brazing or welding, this brazing or welding being able to be carried out intermittently or over the entire length of the plate.
  • Each flank 71 , 72 can optionally have crease lines which define different facets so they can adapt to the different forms of feet 621 , as can be seen in FIGS. 7 and 8 , for example.
  • the plate 7 fulfils both a sealing role by limiting leaks from inter-nozzle sectors 6 and a thermal shielding role with respect to the collar 5 , the abradable material 53 and the other elements of the low-pressure turbine which are below this plate.
  • the plate 8 is arched, as is the shielding plate 7 . It also has a substantially L-shaped transverse section with a fastening flank (not visible in the figure) intended to be fixed by welding or brazing onto the external face 52 of the collar 50 and a shielding flank 81 which extends substantially parallel to the feet 621 .
  • the plate 8 can be sectorised or extend over 360° and have scalloping 810 , or not. Its role is to protect the collar 5 and direct the hot-air flow towards the blades 63 .
  • Each angular sector 6 is mounted on the collar 5 so that its tab 622 is introduced laterally via the open end of the slot 553 , then shifted in an anticlockwise direction, that is, to the left in FIG. 3 .
  • the bottom 555 of the slot 553 plays the role of anti-rotation stop.
  • the rear 541 and front 554 faces constitute support faces for the tab 622 and limit its axial travel and therefore the axial travel of the corresponding angular sector 6 .
  • the tabs 622 are preferably inserted into the slots 553 such that the shoulders 623 or 624 do not make contact with the wing 54 or the arm 552 respectively, effectively leaving slight radial play j 1 , respectively j 2 , between both, and allowing dilation of the different components.
  • FIG. 6 shows an angular nozzle sector 6 the tab of which is inserted into the slot of the coupling carried by the collar 5 .
  • the lug 556 prevents the tab 622 from disengaging, where the elasticity of the arm 552 is sufficient to allow introduction of the tab into the slot.
  • the different angular sectors 6 are all positioned on the external casing of the turbine (see the casing 100 of FIG. 1 ).
  • the collar 5 is brought more closely to the tabs 622 so that each tab 622 is in front of the intake of a slot 553 , and then the collar 5 is turned to the right in FIG. 3 to simultaneously engage all the tabs 622 in the slots 553 .
  • the collar 5 is mounted in the turbine so that its axis X 1 -X′ 1 is joined to the axis X, X′ of the turbine.
  • the collar 5 is cylindrical in form with a fastening system via simple coupling. Its manufacture is therefore not expensive, especially given novel manufacturing processes such as additive manufacturing.
  • the abradable material 53 fixed to the collar 5 is now dissociated from the sectors 6 of the nozzle. Repairs made to the collar 5 or nozzle sectors 6 are therefore independent of each other, and this can simplify maintenance operations.
  • the invention also provides an overall gain in mass over the entire nozzle 4 , since the collar 5 has less thickness than that of the base of an internal platform on a conventional angular sector.
  • the manufacturing cost is also reduced as there is less material and there are fewer zones to control. It is no longer necessary to braze and machine the abradable material on each nozzle sector. Also, there is now no need to make slots on the internal platform 62 , which causes a drop in costs and simplifies manufacturing.
  • the invention overall provides a considerable time gain during installation of the nozzle.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US17/419,404 2018-12-31 2019-12-30 Nozzle for a turbine, turbomachine turbine equipped with said nozzle and turbomachine equipped with said turbine Active US11408295B2 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
FR1874402A FR3091311B1 (fr) 2018-12-31 2018-12-31 Distributeur pour turbine, turbine de turbomachine équipée de ce distributeur et turbomachine équipée de cette turbine.
FR1874402 2018-12-31
PCT/FR2019/053317 WO2020141284A1 (fr) 2018-12-31 2019-12-30 Distributeur pour turbine, turbine de turbomachine équipée de ce distributeur et turbomachine équipée de cette turbine

Publications (2)

Publication Number Publication Date
US20220074312A1 US20220074312A1 (en) 2022-03-10
US11408295B2 true US11408295B2 (en) 2022-08-09

Family

ID=66690613

Family Applications (1)

Application Number Title Priority Date Filing Date
US17/419,404 Active US11408295B2 (en) 2018-12-31 2019-12-30 Nozzle for a turbine, turbomachine turbine equipped with said nozzle and turbomachine equipped with said turbine

Country Status (5)

Country Link
US (1) US11408295B2 (fr)
EP (1) EP3906357B1 (fr)
CN (1) CN113366191B (fr)
FR (1) FR3091311B1 (fr)
WO (1) WO2020141284A1 (fr)

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0017534A1 (fr) * 1979-03-27 1980-10-15 Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." Garniture d'étanchéité amovible pour segment de distributeur de turbomachine
US4239451A (en) * 1978-06-01 1980-12-16 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Device to fasten a seal to the guide vanes of a turbine engine
US6220815B1 (en) * 1999-12-17 2001-04-24 General Electric Company Inter-stage seal retainer and assembly
US20040169122A1 (en) * 2002-10-26 2004-09-02 Dodd Alec G. Seal apparatus
US20130183145A1 (en) 2012-01-17 2013-07-18 Joseph T. Caprario Hybrid inner air seal for gas turbine engines
FR3003599A1 (fr) 2013-03-25 2014-09-26 Snecma Aubage fixe de distribution de flux ameliore
US20150003970A1 (en) 2013-06-27 2015-01-01 MTU Aero Engines AG Sealing arrangement for a turbomachine, a guide vane arrangement, and a turbomachine with such a sealing arrangement
US20180135449A1 (en) * 2016-11-17 2018-05-17 MTU Aero Engines AG Seal system for a guide blade system of a gas turbine
US20180340435A1 (en) 2017-05-29 2018-11-29 MTU Aero Engines AG Seal arrangement for a turbomachine, method for manufacturing a seal arrangement and turbomachine

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP3697705B2 (ja) * 1995-06-22 2005-09-21 石川島播磨重工業株式会社 ガスタービン高圧タービン静翼の固定装置
FR2928961B1 (fr) * 2008-03-19 2015-11-13 Snecma Distributeur sectorise pour une turbomachine.
FR2930592B1 (fr) * 2008-04-24 2010-04-30 Snecma Distributeur de turbine pour une turbomachine
FR2979662B1 (fr) * 2011-09-07 2013-09-27 Snecma Procede de fabrication d'un secteur de distributeur de turbine ou redresseur de compresseur en materiau composite pour turbomachine et turbine ou compresseur incorporant un distributeur ou un redresseur forme de tels secteurs

Patent Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4239451A (en) * 1978-06-01 1980-12-16 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Device to fasten a seal to the guide vanes of a turbine engine
EP0017534A1 (fr) * 1979-03-27 1980-10-15 Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." Garniture d'étanchéité amovible pour segment de distributeur de turbomachine
US6220815B1 (en) * 1999-12-17 2001-04-24 General Electric Company Inter-stage seal retainer and assembly
US20040169122A1 (en) * 2002-10-26 2004-09-02 Dodd Alec G. Seal apparatus
US20130183145A1 (en) 2012-01-17 2013-07-18 Joseph T. Caprario Hybrid inner air seal for gas turbine engines
FR3003599A1 (fr) 2013-03-25 2014-09-26 Snecma Aubage fixe de distribution de flux ameliore
US20150003970A1 (en) 2013-06-27 2015-01-01 MTU Aero Engines AG Sealing arrangement for a turbomachine, a guide vane arrangement, and a turbomachine with such a sealing arrangement
US20180135449A1 (en) * 2016-11-17 2018-05-17 MTU Aero Engines AG Seal system for a guide blade system of a gas turbine
US10677080B2 (en) * 2016-11-17 2020-06-09 MTU Aero Engines AG Seal system for a guide blade system of a gas turbine
US20180340435A1 (en) 2017-05-29 2018-11-29 MTU Aero Engines AG Seal arrangement for a turbomachine, method for manufacturing a seal arrangement and turbomachine

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
French Search Report for FR 1874402 dated Sep. 20, 2019.
International Search Report for PCT/FR2019/053317 dated Mar. 17, 2020 [PCT/ISA/210].

Also Published As

Publication number Publication date
WO2020141284A1 (fr) 2020-07-09
EP3906357A1 (fr) 2021-11-10
FR3091311A1 (fr) 2020-07-03
EP3906357B1 (fr) 2023-03-29
FR3091311B1 (fr) 2021-04-09
CN113366191B (zh) 2023-03-28
CN113366191A (zh) 2021-09-07
US20220074312A1 (en) 2022-03-10

Similar Documents

Publication Publication Date Title
US7217089B2 (en) Gas turbine engine shroud sealing arrangement
US8662835B2 (en) Nozzle for a turbomachine turbine
US8087249B2 (en) Turbine cooling air from a centrifugal compressor
JP4856306B2 (ja) ガスタービンエンジンの流れ通路の静止構成要素
US20100158668A1 (en) Centrifugal compressor forward thrust and turbine cooling apparatus
US8961117B2 (en) Insulating a circumferential rim of an outer casing of a turbine engine from a corresponding ring sector
US20110192166A1 (en) Outlet guide vane structure
US20090104026A1 (en) Control of clearance at blade tips in a high-pressure turbine of a turbine engine
EP3388746A1 (fr) Systèmes de montage de panneau de chambre de combustion et procédés
US20090191050A1 (en) Sealing band having bendable tang with anti-rotation in a turbine and associated methods
CN104797784A (zh) 涡轮机护罩安装和密封结构
US7770401B2 (en) Combustor and component for a combustor
EP2904241B1 (fr) Dispositif anti-erreur de joint d'étanchéité de chambre de combustion de turbine à gaz
US10928067B2 (en) Double skin combustor
JPS62170734A (ja) 遷移ダクトシ−ル構造体
US20220065124A1 (en) Assembly for a turbomachine
US6848885B1 (en) Methods and apparatus for fabricating gas turbine engines
US9528441B2 (en) Aircraft turbofan comprising an intermediate ring with simplified downstream support
US11408295B2 (en) Nozzle for a turbine, turbomachine turbine equipped with said nozzle and turbomachine equipped with said turbine
EP1217231B1 (fr) Assemblage par boulons pour rotors et méthode pour y réduire les gradients thermiques
US11015483B2 (en) High pressure compressor flow path flanges with leak resistant plates for improved compressor efficiency and cyclic life
GB2468848A (en) Turbomachine assembly
US20200386110A1 (en) Sealing structure between turbine rotor disk and interstage disk
JP7271232B2 (ja) 環状燃焼器ライナの遷移ゾーン用の内側冷却シュラウド
WO2021002901A1 (fr) Conduit d'air

Legal Events

Date Code Title Description
AS Assignment

Owner name: SAFRAN AIRCRAFT ENGINES, FRANCE

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:GUERARD, CORALIE CINTHIA;GUESSINE, HAMZA;SIGNING DATES FROM 20191206 TO 20191220;REEL/FRAME:056704/0276

FEPP Fee payment procedure

Free format text: ENTITY STATUS SET TO UNDISCOUNTED (ORIGINAL EVENT CODE: BIG.); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY

STPP Information on status: patent application and granting procedure in general

Free format text: DOCKETED NEW CASE - READY FOR EXAMINATION

STPP Information on status: patent application and granting procedure in general

Free format text: NOTICE OF ALLOWANCE MAILED -- APPLICATION RECEIVED IN OFFICE OF PUBLICATIONS

STCF Information on status: patent grant

Free format text: PATENTED CASE